The present disclosure relates to a gas turbine engine and, more particularly, to a thermal mechanical dimple array for a combustor wall assembly.
Gas turbine engines, such as those that power modern commercial and military aircraft, include a fan section to propel the aircraft, compressor section to pressurize a supply of air from the fan section, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases and generate thrust.
The combustor section may have an annular wall having inner and outer shells that support respective inner and outer heat shielding liners. The liners may be comprised of a plurality of floating heat shields or panels that together define an annular combustion chamber. An annular cooling plenum is defined between the respective shells and liners for supplying cooling air to an opposite hot side of the panels through a plurality of strategically placed cooling or effusion holes. The effusion holes are generally orientated to create a protective blanket or air film over the hot side of the panels, thereby protecting the panels from the hot combustion gases in the chamber.
Unfortunately, and even with an air film, hot spots are still known to exist at isolated locations on each panel. The hotspots create thermal mechanical stresses that lead to cracking and a shortened panel lifespan. Moreover, increasing the density of cooling holes near or at such hotspots can reduce engine efficiency and/or deprive cooling to other locations in the combustor section or in the engine in general.